This invention relates generally to an apparatus for supersonic combustion and more particularly concerns a supersonic combustor for a supersonic combustion ramjet (scramjet) engine.
A ramjet engine is a relatively simple jet-propulsion device in which impinging air is compressed by the forward speed of the aircraft as it enters the engine. Fuel is injected into the compressed air stream and combusted to produce thrust. In a supersonic combustion ramjet, or scramjet, the combustion is carried out in a supersonic air stream. Hydrogen is typically the fuel used with scramjets. The chief advantages of these devices are their simplicity and low weight. A conventional scramjet is shown in cross section in FIG. 1. The device generally comprises three sections: a compression zone 1 where impinging supersonic air (represented by arrow A) enters through an inlet 2 and is compressed, a combustor 3 where the compressed air and fuel injected through a fuel injector 4 are mixed and burned, and an exhaust nozzle 5 where the products of the combustion in the combustor 3 are exhausted to produce forward thrust.
Conventional supersonic aircraft currently operate up to about Mach 3. Their engines are typically turbojets or turbofans in which the air flow through the combustor is slowed to subsonic speeds. This would result in excessively high temperatures for hypersonic flight. Thus, there has been considerable interest in developing a scramjet engine to be used in the propulsion of hypersonic aircraft, i.e., aircraft capable of reaching speeds in the range of Mach 5 to Mach 25. However, for scramjet propulsion to be feasible from an engineering point of view, combustion must be complete within distances on the order of one foot. Thus, the fuel injection must be accomplished in a manner that results in mixing and burning the fuel as rapidly as possible.
The range of conditions confronting aeropropulsion combustors will widen dramatically with the advent of high altitude scramjet engines. For example, while gas-turbine engines might be expected to operate from sea level to about 30 kilometers, corresponding to combustor pressures of 40 atm. to 1-2 atm., scramjet engines envisioned for transatmospheric vehicles will operate at altitudes in excess of 60 kilometers, where ambient pressures are well below 0.1% of atmospheric. Ram compression through inlet/forebody systems of the compression zone can increase the combustor static pressure but only at the expense of mounting aerodynamic losses and heat load on the vehicle. Cooling limitations therefore place an upper bound on the attainable pressures. The concern with low static pressures is that the rate of combustion reactions is lower at lower pressures. Furthermore, the time for combustion is very short in scramjets due to the supersonic flow through the combustor. Even as flight speed drops, the axial velocity through the engine is not significantly reduced. Thus, the "residence time" in a scramjet combustor is typically less than 1.times.10.sup.-4 seconds. The small time duration and slow reaction rate pose obstacles to developing a scramjet combustor which can meet the engineering requirement of completing mixing and combustion within a distance on the order of one foot.
Various scramjet combustors have been previously proposed. One such device is shown in FIG. 1. The fuel injector 4 is situated near the inlet of the combustor 3 and is arranged to provide normal injection of the fuel into the supersonic air flow. As used herein, the term "normal" means perpendicular to the air flow, and the term "axial" means parallel to the air flow. Normal injection is generally beneficial because the injected jet of fuel, being transverse to the air flow, is well-mixed with the compressed air and thus burns more efficiently. However, the normal injection of FIG. 1 has the problem that the pressure rise from the combustion, which occurs forward of the jet, is transmitted upstream of the jet through a subsonic wall boundary layer to the inlet of the combustor. This coupling of the inlet flow and the combustion process creates an aerodynamic disturbance to the inlet flow which will prevent supersonic air flow from continuing into the combustor.
The combustor 3a of a second conventional scramjet is shown in FIG. 2. In this device, the fuel injector 4a provides normal injection of fuel into the supersonic air flow upstream of a backward-facing step 6a formed in a wall of the combustor 3a. This arrangement results in a flame held on the step 6a, downstream of the combustor inlet. However, the injected jet of fuel tends to interact with the inlet, leading to the same coupling problem described above with respect to the device of FIG. 1.
FIG. 3 shows another conventional approach. The fuel injector 4b is arranged in a wall of the combustor 3b so as to produce axial injection of fuel. Axial injection is advantageous in that the fuel has momentum in the axial direction which is added to the overall thrust of the engine. At the high speeds at which scramjets are operated, the structure of the aircraft is greatly heated by friction with the atmosphere. The fuel is typically used to cool the heated structure so the fuel is very hot (approximately 1200.degree.K) upon injection into the compressed air stream. The hot fuel represents a significant source of thrust when injected axially, particularly through a converging or converging-diverging nozzle. The drawback is that axial injection leads to very poor mixing because the fuel and air both flow in the same direction.
In FIG. 4, a strut 7c is disposed in the combustor 3c. The strut 7c has a number of fuel injectors 4c pointing downstream for axial injection. This arrangement is able to take advantage of the increased thrust due to the momentum of the axially injected fuel while achieving good mixing because the fuel is injected in a distributed manner across the air flow. The problem with the strut arrangement is that the strut 7c tends to become very hot and considerable effort must be expended in cooling the strut. Also, aerodynamic losses due to blockage by the strut 7c of the supersonic air flow are significant and decrease overall thrust.
Another previous arrangement is shown in FIG. 5. A fuel injector 4d injects fuel downstream of a backward-facing step 6d formed in a wall of the combustor 3d. The fuel injection can be normal (as shown) or axial (not shown). This arrangement isolates the inlet to the combustor 3d from the combustion process by interrupting the boundary layer, thereby avoiding the problem of coupling between the inlet flow and the combustion process present in the arrangements of FIGS. 1 and 2. However, computational and experimental studies (see Correa, S. M. and Warren, R. E. "Supersonic Sudden-Expansion Flow with Fluid Injection: an Experimental and Computational Study," Paper AIAA 89-389, 1989) have shown that with this arrangement the flow behind the step 6d does not penetrate into the main flow, thus limiting mixing. This increases the combustor length needed to achieve adequate mixing.
Thus, there remains a need for a supersonic scramjet combustor capable of completing mixing and combustion within a distance on the order of one foot while using axial and/or normal fuel injection.